The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine, air is pressurized in a multistage compressor and mixed with fuel for generating hot combustion gases in a combustor. The gases are discharged through a high pressure turbine (HPT) which powers the compressor, typically followed by a low pressure turbine (LPT) which provides output power by typically powering a fan at the upstream end of the engine. This turbofan configuration is used for powering commercial or military aircraft.
Engine performance or efficiency may be increased by increasing the maximum allowed operating temperature of the combustion gases that are discharged to the HPT which extracts energy therefrom. Furthermore, engines are continually being developed for increasing cruise duration and distance, for one exemplary commercial application for a supersonic business jet and for an exemplary military application such as a long range strike aircraft.
Increasing turbine inlet temperature and cruise duration correspondingly increases the cooling requirements for the hot engine components, such as the high pressure turbine rotor blades. The first stage rotor blades receive the hottest combustion gases from the combustor and are presently manufactured with state-of-the-art superalloy materials having enhanced strength and durability at elevated temperature. These blades may be configured from a myriad of different cooling features for differently cooling the various portions of the blades against the corresponding differences in heat loads thereto during operation.
The presently known cooling configurations for first stage turbine blades presently limit the maximum allowed turbine inlet temperature for obtaining a suitable useful life of the blades. Correspondingly, the superalloy blades are typically manufactured as directionally solidified materials or monocrystal materials for maximizing the strength and life capability thereof under the hostile hot temperature environment in the gas turbine engine.
The intricate cooling configurations found in the blades are typically manufactured using common casting techniques in which one or more ceramic cores are utilized. The complexity of the cooling circuits in the rotor blades is limited by the ability of conventional casting processes in order to achieve suitable yield in blade casting for maintaining competitive costs.
Like the first stage turbine blades, the first stage turbine nozzle includes hollow vanes which require suitable cooling for extended life while exposed to the hot combustion gases. The vanes, like the blades have corresponding airfoil configurations, and include internal cooling circuits of various configurations specifically tailored to cool the different parts of the vanes corresponding with the different heat loads from the combustion gases.
Accordingly, it is desired to provide a turbine airfoil having an improved cooling configuration for further advancing temperature and durability thereof in a gas turbine engine.